To increase the efficiency of gas turbine engines, a known approach is to raise the turbine operating temperature. As operating temperatures are increased, the thermal limits of certain engine components may be exceeded, resulting in material failure or, at the very least, reduced service life. In addition, the increased thermal expansion and contraction of these components adversely effects clearances and their interfitting relationships with other components of different thermal coefficients of expansion. Consequently, these components must be cooled to avoid potentially damaging consequences at elevated operating temperatures. It is common practice then to extract from the main airstream a portion of the compressed air at the output of the compressor for cooling purposes. So as not to unduly compromise the gain in engine operating efficiency achieved through higher operating temperatures, the amount of extracted cooling air should be held to a small percentage of the total main airstream. This requires that the cooling air be utilized with utmost efficiency in maintaining the temperatures of these components within safe limits.
A particularly critical component subjected to extremely high temperatures is the shroud located immediately beyond the high pressure turbine nozzle from the combustor. The shroud closely surrounds the rotor of the high pressure turbine and thus defines the outer boundary of the extremely high temperature, energized gas stream flowing through the high pressure turbine. To prevent material failure and to maintain proper clearance with the rotor blades of the high pressure turbine, adequate shroud cooling is a critical concern.
One approach to shroud cooling, such as disclosed in commonly assigned U.S. Pat. Nos. 4,303,371--Eckert and 4,573,865--Hsia et al., is to provide various arrangements of baffles having perforations through which cooling air streams are directed against the back or radially outer surface of the shroud to achieve impingement cooling thereof. Impingement cooling, to be effective, requires a relatively large amount of cooling air, and thus engine efficiency is reduced proportionately.
Another approach is to direct a film of cooling air over the front or radially inner surface of the shroud to achieve film cooling thereof. Unfortunately, the cooling air film is continuously being swept away by the spinning rotor blades, thus diminishing film cooling effects on the shroud.
It is accordingly an object of the present invention to provide an improved cooling assembly for maintaining the shroud in the high pressure turbine section of a gas turbine engine within safe temperature limits.
A further object is to provide a shroud cooling assembly of the above-character, wherein effective shroud cooling is achieved using a lesser amount of pressurized cooling air.
An additional object is to provide a shroud cooling assembly of the above-character, wherein the same cooling air is applied in a succession of cooling modes to maximize shroud cooling efficiency.
Another object is to provide a shroud cooling assembly of the above-character, wherein heat conduction from the shroud into the supporting structure therefor is reduced.
Other objects of the invention will in part be obvious and in part appear hereinafter.